1. Field of the Invention
The present invention relates to a working method of a cooling air passage for the flow of cooling air in a gas turbine stationary blade shroud.
2. Description of the Prior Art
As a high temperature gas turbine is being developed, the cooling of a shroud of gas turbine stationary blade is also becoming necessary. FIG. 2 shows an entire shroud and FIG. 3 shows a cross sectional view, taken on line 3--3 of FIG. 2, of a cooling air passage in the prior art provided in the shroud.
In FIG. 3, a cooling air passage 11 is provided, passing through a shroud 12 in an axial direction, below a seal groove 13 in which adjacent shrouds 12 fit with each other. As shown in FIG. 2, the cooling air passage 11 is very long as compared with its diameter. Therefore, in the working of the cooling air passage 11, a high grade working technology is required.
So, in order to drill the cooling air passage 11 long enough relative to its diameter, machining work, keeping away from the nearby seal groove 13, becomes necessary. But due to the length of the passage, machining work with precise accuracy is very difficult.
Further, in case a turbulator (cooling fin) is to be provided in the cooling air passage, working with accuracy will be almost impossible.